This invention described herein is a method for orienting spacecraft, and more specifically, is a novel method for maintaining a satellite in a synchronous orbit in a way which uses substantially less fuel than conventional methods.
With reference to FIG. 1, it is known that a satellite 2 orbiting the earth 1 at a distance A of approximately 6.61 times the radius R.sub.E of the earth (one earth radius R.sub.E is about 3964 miles or 6378 kilometers) from the center 3 of the earth will orbit the earth in a sidereal day, i.e. the period of time for the earth to complete its rotation (23 hours, 56 minutes and 4 seconds). Therefore, a satellite 2 in such an orbit is known as a geosynchronous satellite. When the orbit of the satellite 2 is in a plane 5 encompassing the equator 4 (equatorial plane 5), the satellite, as viewed from a subsatellite point E.sub.S on the equator, will appear stationary. Therefore, a satellite in a geosynchronous orbit in the equatorial plane is referred to as a geostationary satellite.
In actuality, as shown in FIG. 2, the satellite 2 in a geostationary orbit will experience departures from an ideal orbital path P. The departures are caused mainly by attraction forces such as the gravitational pulls of the sun and moon, and the oblateness of the earth. Due to the departures, the satellite 2 will appear to be moving in a small cyclical pattern during any given sidereal day, as viewed from the subsatellite point E.sub.S. Movements of the satellite above the equatorial plane 5 are referred to as excursions to the north N, and movements below the equatorial plane 5 are referred to as excursions to the south S. Movements to the left or right of the point on the orbital path P at which the satellite is located are referred to as excursions to the east E or the west W, respectively. If these excursions are prevented, then the satellite 2 will appear stationary to a viewer located at the subsatellite point E.sub.S, as previously mentioned.
As is well-known in the art, the satellite 2 may be stabilized by rotational inertia about a spin axis or angular momentum axis 13 (hereinafter, spin axis includes angular momentum axis), as shown in FIG. 1. Satellites are stabilized in orbit by means of an internal gyroscope, a rotating portion of structure, firing thrusters or a combination thereof. As used herein, satellites stabilized by rotating a portion of their structure, such as a Hughes type 376 satellite, are called spin stabilized. These satellites also employ thrusters for attitude control. The spin axis 13 of the satellite 2 in a geostationary orbit is usually kept at equatorial-normal E.sub.N (perpendicular to the equatorial plane 5). However, due to forces such as "solar pressure" (photons striking the satellite), the spin axis 13 will tend to precess. Controlling the amount of precession is referred to as spin axis attitude control.
In communications satellites and the like, it is conventional to use geostationary satellites so that a generally conical beam 12 from an antenna 11 on the satellite 2 can continuously cover a specified region of the earth, e.g. the continental U.S. In FIG. 1, to aid in the understanding of the geometry involved in methods of orienting geostationary satellites, the target region is illustrated as a relatively small region 6 and the equatorial orbital plane 5 is depicted as intersecting the antenna 11 rather than the center of mass of the satellite. However, if the position of the satellite 2 is not controlled, the satellite initially in a geostationary orbit will experience the above-mentioned departures. For example, the plane in which the satellite 2 orbits (orbital plane) will incline with respect to the equatorial plane 5 at a rate between 0.75.degree. to 0.95.degree. per year. East/west excursions as well as precessions or attitude changes of the spin axis 13 will also occur. As a result, the axis 112 of the beam 12 will be shifted from a bore site target 66. The shifting of the beam axis 112 can become sufficiently large so that the beam 12 will not adequately cover the target site 6. Therefore, to keep the satellite 2 stationary so that the beam 12 covers the target site 6 and the beam axis 112 points at the bore site target 66, operations known as stationkeeping and spin axis control are performed. That is, the north/south excursions, east/west excursions and the attitude of the spin axis 13 are controlled in a well-known manner, so that the satellite 2 is kept substantially in the equatorial plane 5 with the spin axis at a desired attitude. Therefore, the cyclical "figure eight" pattern shown in FIG. 2 is kept very small, and the beam 12 is kept essentially stationary on the target site 6.
Currently, satellites are provided with thrusters and fuel (or other propellants) for performing stationkeeping. As is well-known in the art, the necessary thrusting parameters (e.g. thruster on time and phasing or frequency of thrusting) to achieve a desired orbital position are determined using an orbital mechanics computer program. By commanding the satellite 2 to implement the calculated thrusting parameters, the satellite may be kept substantially in the desired geostationary orbit.
Stationkeeping in accordance with current technique requires a large amount of fuel for the thrusters. Since fuel is heavy and weight limitations are critical considerations in the design of a satellite, it is common that the limit on a geostationary satellite's useful life is determined by the amount of on board fuel available for the thrusters. For example, a conventional satellite such as COMSTAR (Trademark), uses an average of 37 pounds of fuel for stationkeeping during each year of the latter part of its design life. Of the 37 pounds of fuel consumed in a year, approximately 34 pounds are used for north/south correction, while only 2 pounds are used for east/west correction and 1 pound for attitude control. Since such a conventional satellite is provided with approximately 340 pounds of fuel and uses more of that fuel in the earlier part of its design life, it is expected to run out of fuel in just over 7 years. When a satellite is out of fuel, the excursions and tilting can no longer be controlled. Therefore, the antenna 11 cannot be kept continuously pointed at the target site 6 and the satellite's usefulness ends.
Currently, only a limited amount of fuel in stored on board, as the design weight of satellites is strictly controlled. Due to this limited supply of fuel, the useful life of a satellite is normally determined by the amount of fuel stored on board. Thus, the prior art suffers from the disadvantage that a substantial portion of a satellite's weight must be fuel, and the useful life of the satellite is severely limited by its fuel capacity. Moreover, the satellite must be launched substantially into the equatorial plane. If the satellite is not launched accurately, fuel must be expended to move the satellite substantially into the equatorial plane.
A proposed method for keeping the antenna of a body-stabilized satellite pointed at the target site without north/south correction is disclosed in U.S. Pat. No. 4,084,772 (hereinafter, "the '772 patent). In the proposed method, the body-stabilized satellite has a pitch axis of a type automatically kept along a line normal to the orbital plane (orbit-normal) by a closed loop roll control system. The method further involves equipping the satellite with a transverse momentum wheel mounted with its momentum axis parallel to the yaw axis, and varying the transverse momentum vector in a sinusoidal pattern having a period equal to the orbital period in order to vary the position of the pitch axis with respect to orbit-normal and point the satellite's antenna at the bore site target. The proposed method thus involves corrections being made continuously during each orbit, and using a satellite equipped with a closed loop roll correcting system. However, equipping a satellite with a closed loop roll correcting system necessitates increasing the weight of the satellite. Moreover, spin stabilized satellites generally do not have such a closed loop control system, and it would be expensive to provide a spin stabilized satellite with such a system. In addition, the method of the '772 patent cannot be used in satellites lacking a transverse momentum wheel. Thus, the method cannot be used on many satellites currently in orbit. Further, to design and build satellites with a transverse momentum wheel will increase the weight of the satellite and increase the cost thereof.